The present invention relates generally to gas turbine engines, and, more specifically, to turbine shrouds and blade tips therein.
In a gas turbine engine, air is pressurized in a compressor, mixed with fuel in a combustor, and ignited for generating hot combustion gases which flow downstream through one or more turbine stages which extract energy therefrom. A high pressure turbine (HPT) first receives the combustion gases from the combustor and extracts energy therefrom for powering the compressor. A low pressure turbine (LPT) follows the HPT for extracting additional energy for providing output energy typically used for powering a fan disposed upstream of the compressor in a typical aircraft gas turbine engine application.
The HPT includes a stationary turbine nozzle having a plurality of circumferentially spaced apart stator vanes which control discharge of combustion gases from the combustor. The HPT also includes at least one rotor stage having a plurality of circumferentially spaced apart turbine rotor blades extending radially outwardly from a supporting rotor disk. The blades include airfoils which receive combustion gases from the nozzle and extract energy therefrom for rotating the rotor disk and in turn rotating the compressor. The airfoils are typically hollow and include internal cooling circuits therein through which a portion of pressurized air bled from the compressor is channeled for cooling the blades.
Surrounding the rotor blades is an annular turbine shroud fixedly joined to a surrounding stator casing. The shroud is suspended closely atop the blade tips for providing a small gap or tip clearance therebetween. The tip clearance should be as small as possible to provide an effective fluid seal thereat during operation for minimizing the amount of combustion gas leakage therethrough for maximizing efficiency of operation of the engine. However, due to differential thermal expansion and contraction of the rotor blades and surrounding turbine shroud, the blade tips occasionally rub against the inner surface of the shroud causing abrasion wear.
Since the blade tips are at the radially outermost end of the rotor blades and are directly exposed to the hot combustion gases, they are difficult to cool and the life of the blade is thereby limited. The blade tips are typically in the form of squealer rib extensions of the pressure and suction sides of the airfoil, extending outwardly from a tip floor. Cooling air is channeled under the floor to cool the ribs by conduction. And, film cooling holes may extend through the floor to film cool the exposed ribs.
Since the turbine shroud is also exposed to the hot combustion gases, it too is also cooled by bleeding a portion of the pressurized air from the compressor, which is typically channeled in impingement cooling against the radially outer surface of the turbine shroud. Turbine shrouds typically also include film cooling holes extending radially therethrough with outlets on the radially inner surface of the shroud from which is discharged the cooling air in a film for cooling the inner surface of the shroud.
The holes are typically arranged in a pattern between the forward and aft axial ends of the shroud to provide uniform discharge of the cooling air through the shroud. Cooling air used to cool the blade tips and turbine shroud has limited effectiveness, and decreases the overall efficiency of the engine.
Accordingly, it is desired to provide an improved turbine shroud for cooperating with turbine rotor blade tips for improving cooling of the shroud, as well as the blade tips.